The present invention relates to a composite structure element with built-in damping capacity which will minimize and, in some instances, eliminate a need for expensive active structural-acoustic control systems, as well as active control non-acoustic disturbance control systems in spacecraft, aircraft structures, automobiles, precision machinery and the like. The invention further relates to apparatus utilizing the composite structure element and to a method of making the same.
During launch of a spacecraft, the composite structures including composite payload fairings of the spacecraft are subject to a broad band of frequencies, high level acceleration disturbances. All composite components of a spacecraft do not have sufficient damping capabilities to suppress these vibrations and, therefore require spacecraft structure to be actively controlled. For example, commonly assigned U.S. Pat. No. 5,655,757 discloses an actively controlled damper which resists relative vibration of parts, and is particularly directed to an actively controlled damper which has a ring-shaped structure for damping vibration between a payload and an adapter fitting of a rocket launch vehicle.
Another problem arises in achieving fast settling times of a spacecraft. An undamped system with a very low natural frequency requires a long settling time during slew of the spacecraft which is undesirable.
An object of the present invention is to provide an improved composite structure element with built-in damping capacity which will minimize and in some instances eliminate the need for expensive active structural-acoustic control systems and active control non-acoustic disturbance control systems. The composite structure element of the invention can advantageously be used in spacecraft and also in aircraft, automobiles and in precision machinery.
A composite structure element with built-in damping according to one form of the invention comprises a non-metallic matrix, a plurality of filaments within the matrix reinforcing the matrix, and a lightweight metal having a specific damping capability of at least 1% provided on at least one of the matrix and the filaments of the composite structure element for attenuating vibrations of the element. In the disclosed embodiments, the lightweight metal is preferably selected from the group consisting of a magnesium alloy and an aluminum alloy. Most preferably, a magnesium alloy containing 0.1-10 wt. % zirconium, balance essentially magnesium is utilized for obtaining excellent damping capacity.
The composite structure element according to a first embodiment of the invention comprises a matrix of plastic about a plurality of filaments which are coated with a lightweight metal having a specific damping capability of at least 1%. The thickness of the coating on the filaments is preferably between 0.1 and 100 times a diameter of the filaments.
The composite structure element of a second embodiment of the invention comprises a porous, non-metallic matrix with the pores thereof being filled with a lightweight metal having a specific damping capability of at least 1%. In a disclosed example, the porous matrix is a carbonaceous material which is reinforced with filaments, e.g., carbon fibers, which may optionally be coated with a lightweight metal having a specific damping capability of at least 1%.
The filaments of the composite structure element can be continuous in length, extending unidirectionally as in a wound structure, or the filaments can extend bidirectionally, as with the use of a woven cloth of filaments. As another variation, the filaments can be relatively short and have a random directional orientation in the matrix. The filaments are preferably carbon fibers embedded in the non-metallic matrix in the disclosed embodiments, but other reinforcing fibers such as glass and aramid fibers could be used.
A disclosed apparatus of the invention utilizing the composite structure element comprises a rocket launch vehicle including an adapter interface fitting, a rocket payload mounted on the rocket launch vehicle by way of the adapter interface fitting, and a damper located between the rocket payload and the adapter interface fitting for damping vibrations between the rocket launch vehicle and the rocket payload. The damper includes at least one composite structure element according to the invention for attenuating vibrations.
The present invention further comprises a fairing for a spacecraft, particularly a payload fairing of a rocket, a spacecraft bus and a damper element of a rocket, formed with the composite structure elements of the invention for providing these components with a high specific damping to minimize or, in some instances, eliminate the need for active structural-acoustic control systems and active control non-acoustic disturbance control systems.
The composite structure element with built-in damping in a further embodiment of the invention comprises a metal mesh and a lightweight metal having a specific damping capability of at least 1% provided on the metal mesh for attenuating vibrations of the element. In the disclosed embodiment, the metal mesh is wire mesh of a satellite antenna on which a lightweight metal having a specific damping capability of at least 1% and selected from the group consisting of a magnesium alloy and an aluminum alloy, is provided for attenuating vibrations of the antenna. A related method of the invention for making a composite structure element with built-in damping comprises the steps of providing a structural member which is subject to being disturbed when used for its intended purpose and providing a lightweight metal having a specific damping capability of at least 1% on the structural member for damping disturbances of the structural member during its use.
These and other objects, features and advantages of the present invention will become more apparent from the following description when taken in connection with the accompanying drawings, which show, for purposes of illustration only, several embodiments in accordance with the present invention.